Gas turbine high temperature turbine blade outer air seal assembly

ABSTRACT

A turbine shroud assembly includes forward and aft hangers, an axisymmetric plenum assembly, ceramic shroud segments, ceramic spacers, and forward and aft rope seals. The plenum assembly supplies impingement cooling to the shroud and the hangers. The impingement cooling to the forward and aft hangers is controlled independently to improve blade tip clearance. The rope seals are radially inward from the hangers and reduce cooling flow leakage. The turbine shroud assembly can operate in a higher temperature environment using less cooling flow than the prior art.

GOVERNMENT INTERESTS

The invention was made with Government support under contracts with theUS Army (DAAJ02-94-C-0030), the US Air Force (F33615-94-C-2507) and theUS Navy (N00019-94-C-0036). The Government has certain rights in thisinvention.

BACKGROUND OF THE INVENTION

The present invention generally relates to gas turbine engine systemsand, more particularly, to high pressure turbine (HPT) blade outer airseal (BOAS) assemblies, also known as turbine shroud assemblies.

Turbine shroud assemblies have been used extensively in gas turbineengines. The turbine shroud assembly may be positioned immediatelydownstream of an HPT nozzle. The turbine shroud assembly may surround aHPT rotor and may define an outer boundary of a high temperature gasflow path through the HPT. During engine operation, exposure to the hightemperature gas flow may result in failure of the turbine shroudcomponents. Due to the differing expansion of rotor and turbine shroudassembly components, it may also result in contact between the turbineshroud assembly and the blade tips of the rotor. A small amount ofcooling air from a compressor may be used to decrease some of theadverse effects of the high temperature gas flow.

Minimizing the amount of air necessary to cool the turbine shroudassembly is desirable because engine efficiency decreases as the amountof cooling air increases. Methods for minimizing the cooling airnecessary may include decreasing cooling air leakage from the assemblyor reducing the cooling needs of the system by increasing theeffectiveness of the cooling scheme.

Turbine shroud assemblies have experienced significant distress due to alack of robust sealing of the assembly. This leakage may result in asignificant reduction in the cooling cavity pressure (and back flowmargin), which can result in hot gas ingestion and distress in thehardware. Back flow margin is the ratio of the difference between theshroud cooling cavity pressure and the flow path pressure to the flowpath pressure. If the back flow margin of the assembly becomes negative(or for some designs even a low positive number), hot flow path gas mayingest into portions of the shroud and may cause significant distress.The challenge in maintaining good back flow margin is due to thedifficulty in sealing the various leak paths that allow the cooling airto escape from the shroud cooling cavity.

Several methods of reducing cooling air leakage have been disclosed.These methods include the use of labyrinth type seals and metallicplatform seals. Unfortunately, labyrinth seals are not suitable for someapplications, and the metallic platform seals, which are secured inmachined grooves in the sides of the segments, may fail in the operatingenvironment of some engines. In addition, assembly technicians may cutthemselves on the small, sharp metallic platform seals.

Methods of reducing system cooling needs have also been disclosed.Manufacturing the assembly components from more robust materials andutilizing thermal barrier coatings (TBC) have been described. Designsthat utilize TBC to keep the shrouds insulated from the hot flow pathgas can experience delamination of the TBC, which may result in shrouddistress, which may result in large turbine blade tip clearances. Thesubsequent increase in turbine blade tip clearance significantly hurtsfuel consumption and also results in an increase in turbine inlettemperature, which further distresses the hardware.

Methods of increasing the effectiveness of cooling configurations havebeen disclosed. In one method complex arrays of film cooling holes havebeen drilled into shroud segments. Although this results in increasedcooling of the turbine shroud assembly, all edges of the shroud segmentsmay not be sufficiently cooled and system integrity may suffer.

Turbine shroud assemblies having increased cooling of the shroud segmentedges have been disclosed in U.S. Pat. No. 6,270,311. This inventionutilizes an interlocking hook/shelf on the ends of the segments inconjunction with conventional feather seals and slots to produce an endgap seal between the adjacent circumferential segments. In addition,this invention uses film cooling holes to reinforce cooling at the sidesof the segment. Although cooling of the shroud segment edges isincreased, the metallic feather seals may suffer distress at higheroperating temperatures, which may result in a loss of back flow marginto the assembly.

Another turbine shroud assembly has been disclosed in U.S. patentapplication No. 2003/0133790. This invention requires that the turbineshroud segment and the shroud segment hanger both are segmented arcs.This invention relies on tight tolerances to minimize leakage of theassembly. Sliding the turbine shroud segment into the shroud segmenthanger requires tight tolerances to keep the air seal along the forwardand aft hooks. Unfortunately, the tolerances needed may result inincreased production costs of the turbine shroud assembly. In addition,fine-tuning of the thermal expansion of the forward and aft hangers maynot be possible. Further, to change cooling flows, this inventionrequires that every shroud segment hanger be reworked.

Turbine shroud assemblies having increased cooling efficiency have beendisclosed in U.S. Pat. No. 5,188,506. This design incorporates a ropeseal radially outside of the segment forward hook to reduce the leakageof cooling air through the forward hook region of the shroud support.Unfortunately, the forward hook of this assembly may be exposed to hotingested flow path air. Also, fine tuning of the thermal growth of theforward hanger may not be possible. Further, the disclosed assembly maynot allow for sufficient axial motion of the shroud segment as thepressure loads move the segment aft. Because turbine inlet temperatureswill continue to rise to achieve greater thrust to weight capability andimproved fuel consumption, still further improvements are needed.

As can be seen, there is a need for improved turbine shroud assemblies.Additionally, assemblies are needed wherein cooling air flow isminimized while allowing for increased gas flow temperatures. Further,turbine shroud assemblies are needed wherein blade tip clearance isdecreased. Moreover, turbine shroud assemblies having improved coolingschemes to the forward and aft hangers are needed. Also, assemblies areneeded that have reduced cooling air leakage and improved shroud segmentsealing.

SUMMARY OF THE INVENTION

In one aspect of the present invention, a turbine shroud assemblycomprises a plurality of shroud segments assembled circumferentiallyabout a longitudinal engine centerline axis, each shroud segment havinga forward hook and an aft hook; a plurality of spacer channelspositioned on the radially outward side of the shroud segments, suchthat one spacer channel is in contact with each interface of two shroudsegments; a plurality of ceramic spacer seals positioned in contact withthe spacer channels, such that one ceramic spacer seal is within eachspacer channel; a forward hanger, comprised of a plurality of forwardhanger rails positioned radially outward from the shroud segments, theforward hanger rails capable of engaging the shroud segment forwardhooks, the forward hanger rails having an o-ring groove positionedcircumferentially on a radially inward side; an aft hanger positionedradially outward from the shroud segments, the aft hanger comprised of arail capable of engaging the shroud segment aft hooks, the aft hangerrail having an angled surface positioned on the forward edge of theradially inward side; and a plenum assembly positioned between and incontact with the forward hanger and the aft hanger.

In yet another aspect of the present invention, a turbine shroudassembly comprises a plurality of ceramic shroud segments assembledcircumferentially about a longitudinal engine centerline axis, eachceramic shroud segment having a forward hook and an aft hook; and aplurality of ceramic spacer seals positioned in contact with the ceramicshroud segments, such that each ceramic spacer seal is in contact withthe radially outward side of two ceramic shroud segments.

In another aspect of the present invention, a turbine shroud assemblycomprises a plurality of shroud segments assembled circumferentiallyabout a longitudinal engine centerline axis, each shroud segmentcomprising a monolithic silicon nitride ceramic and having a forwardhook and an aft hook; a plurality of spacer channels positioned on theradially outward side of the shroud segments, such that one spacerchannel is in contact with each interface of two shroud segments; aplurality of ceramic spacer seals comprising a monolithic siliconnitride ceramic, the ceramic spacer seals positioned in contact with thespacer channels, such that one ceramic spacer seal is within each spacerchannel; a forward hanger comprising a nickel based alloy, positionedradially outward from the shroud segments, the forward hanger railscapable of engaging the shroud segment forward hooks, the forward hangerrails having an o-ring groove positioned circumferentially on a radiallyinward side; a forward rope seal positioned between and in contact withthe o-ring groove and the shroud segments; an aft hanger comprising anickel based alloy, positioned radially outward from the shroudsegments, the aft hanger rail capable of engaging the shroud segment afthooks, the aft hanger rail having an angled surface positioned on theforward edge of the radially inward side; an aft rope seal positionedbetween and in contact with the angled surface and the shroud segments;a plenum assembly positioned between and in contact with the forwardhanger and the aft hanger, the plenum assembly comprising anaxisymmetric plenum balloon having an impingement cooling array therethrough, a plurality of flow metering openings in fluid communicationwith the axisymmetric plenum balloon, and a plurality of inlet openingsin flow communication with the flow metering openings.

In another aspect of the present invention, an apparatus for a turbineengine comprises an axisymmetric plenum balloon having an impingementcooling array there through; a plurality of flow metering openings influid communication with the axisymmetric plenum balloon; and aplurality of inlet openings in flow communication with the flow meteringopenings.

In still another aspect of the present invention, a rope seal apparatusfor use between a turbine shroud and a turbine hanger comprises acompressed hybrid ceramic rope positioned between and in contact withthe turbine shroud and the turbine hanger, such that the turbine hangeris radially outward from the compressed hybrid ceramic rope.

In a further aspect of the present invention, a method of shielding aturbine engine structure from a hot gas flow path there throughcomprises the steps of providing a plurality of ceramic shroud segmentsassembled circumferentially about a longitudinal engine centerline, aforward hanger radially outward from and forward of the ceramic shroudsegments, an aft hanger radially outward from and aft of the ceramicshroud segments, and a plenum assembly between and in contact with theforward hanger and the aft hanger; and supplying a cooling flow to theplenum assembly such that the cooling flow impinges the ceramic shroudsegments, the forward hanger, and the aft hanger.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdrawings, description and claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial, cut-away perspective view of a turbine shroudassembly according to one embodiment of the present invention;

FIG. 2 is an exploded, partial view of FIG. 1 according to oneembodiment of the present invention;

FIG. 3 is a cross sectional view of a turbine shroud assembly accordingto one embodiment of the present invention;

FIG. 4 is a cross sectional view of a plenum assembly according to oneembodiment of the present invention;

FIG. 5 is a partial sectional view looking aft of a turbine shroudassembly according to one embodiment of the present invention;

FIG. 6 is a partial cross sectional view of a turbine shroud assemblyaccording to one embodiment of the present invention;

FIG. 7 is a partial sectional view looking forward of a turbine shroudassembly according to one embodiment of the present invention; and

FIG. 8 is a partial cross sectional view of an aft rope seal accordingto one embodiment of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

The following detailed description is of the best currently contemplatedmodes of carrying out the invention. The description is not to be takenin a limiting sense, but is made merely for the purpose of illustratingthe general principles of the invention, since the scope of theinvention is best defined by the appended claims.

The present invention generally provides high pressure turbine (HPT)blade outer air seal (BOAS) assemblies, also known as turbine shroudassemblies and methods for producing the same. The turbine shroudassemblies produced according to the present invention may findbeneficial use in many industries including aerospace and industrial.The turbine shroud assemblies of the present invention may be beneficialin applications including electricity generation, naval propulsion,pumping sets for gas and oil transmission, aircraft propulsion,automobile engines, and stationary power plants. This invention may beuseful in any gas turbine application.

In one embodiment, the present invention provides a turbine shroudassembly for an HPT stage. The turbine shroud assembly may be positionedimmediately downstream of an HPT nozzle. The turbine shroud assembly maysurround an HPT rotor and may define an outer boundary of a hightemperature gas flow path through the HPT. Cooling flow from acompressor may be utilized to cool the turbine shroud assembly. Theturbine shroud assembly of the present invention may comprise a plenumassembly positioned between forward and aft hangers. Unlike the priorart, the plenum assembly may comprise an axisymmetric plenum balloonhaving an impingement cooling array there through. The impingementcooling array may provide cooling flow to the shroud segments and,unlike the prior art, it may also provide cooling flow to the forwardand aft hangers. One of the significant advantages of this design overthe prior art may be the ability to customize the heat transfer to theforward and aft hangers, which govern the radial position of the shroudsegments, which in turn govern the turbine rotor tip clearance. Thisunique advantage over the prior art helps maintain tight turbine tipclearance for improved performance. Another advantage over the prior artis that any leakage from the combustor plenum that occurs between theforward hanger and the plenum assembly, or from between the aft hangerand the plenum assembly, may not be wasted air. It may act tosupercharge the impingement cooling air cavity of the turbine shroudassembly to help maintain positive back flow margin to the assembly.

The plenum assembly of the present invention, unlike the prior art, maycomprise inlet openings and flow metering openings. These openings mayprovide a method of pressure recovery for axial flow combustor plenums,may serve to minimize the cooling flow circuit sensitivity to varying oruncertain orifice discharge coefficients that are common with combustorplenums that suffer from significant circumferential swirl, maystraighten the cooling flow as it flows radially inward, and may controlthe overall amount of cooling flow to the plenum assembly. Further,unlike the prior art, a cooling flow change can be made to the presentturbine shroud assembly by simply modifying the flow metering openingdiameter or number of flow metering openings. With the prior artdesigns, each of the shroud segments may be required to undergomodifications. In addition, many engine designs require that the turbinecase undergo modifications as well. This may not be required with thepresent invention.

The turbine shroud assembly of the present invention may comprise hybridceramic rope seals at the forward and aft portions of the shroudassembly. Unlike the prior art, the hybrid ceramic rope seals of thepresent invention may be positioned radially inward from the hangers toreduce hanger exposure to the hot gas flow. Cooling flow leakage may bereduced and cooling cavity pressure may be maintained. By controllingthe relative compression in the hybrid ceramic rope seals, an optimizedflow distribution may be maintained through the turbine shroud assemblythat is not possible in the prior art. Additionally, unlike the priorart that allows the air to escape far from the flow path, the coolingflow from the hybrid rope seals may be channeled directly along theshroud segment to help cool the shroud segment and provide purging ofthe hot flow path gas away from the turbine shroud assembly.

As seen in FIG. 1, a turbine shroud assembly 25 may comprise a shroudsegment 30, a ceramic spacer seal 38, a forward hanger 32, an aft hanger33, a vertical flange 39, and a plenum assembly 40. As better seen inFIG. 2, a turbine shroud assembly 25 may further comprise a forward ropeseal 34 and an aft rope seal 35. A turbine shroud assembly 25 mayfurther comprise a thermal barrier coating (TBC) 31, a forward spacerradial retainer 36, and an aft spacer radial retainer 37, as seen inFIG. 3.

The turbine shroud assembly 25 may comprise a plurality of shroudsegments 30. Each shroud segment 30 may comprise a forward hook 54 andan aft hook 55 capable of engaging the forward hanger rails 56 and afthanger rail 58, respectively. The shroud segments 30 may be assembledcircumferentially about a longitudinal engine centerline axis 71, asshown in FIG. 1. The shroud segments 30 are positioned such that theysurround a HPT rotor (not shown) and define an outer boundary of a hightemperature gas flow path through the HPT. The shroud segments 30 maycomprise a very high temperature capable ceramic, such as monolithicsilicon nitride ceramic. Useful ceramics may include AS800® siliconnitride monolithic ceramic available from Honeywell International, Inc.The shroud segments 30 may be machined out of a full ring and cut intosegments. The shroud segments 30 may have a thickness between about 0.25and about 0.50 inches.

Each shroud segment 30 may be provided with two edge channels 67 a and67 b, one on each circumferential end of the shroud segment. As shown inFIG. 5, each edge channel 67 a and 67 b is comprised of a flat radialsurface 53, a flat circumferential surface 52, and a fillet 68positioned at the interface of flat radial surface 53 and flatcircumferential surface 52. The flat circumferential surface 52 includesthe edge of a shroud segment 30 that is adjacent to the edge of a secondshroud segment 30 when the shroud segments 30 are positionedcircumferentially in the shroud assembly. Two edge channels 67 a and 67b may be positioned at a platform gap 62 between two circumferentiallyassembled shroud segments 30 to form a spacer channel 67. The spacerchannel 67 may be capable of receiving a ceramic spacer seal 38 tocreate seal 66 at the interface between flat circumferential surface 52and the radial inward surface of the ceramic spacer seal 38, thusminimizing cooling flow leakage out the platform gap 62. The edgechannels 67 a and 67 b may have a depth between about 0.05 and about0.40 inches. The edge channels 67 a and 67 b may be provided by knownmanufacturing methods. The shroud segments 30 may be thermallycontrolled by utilizing a thermal barrier coating (TBC) 31 on the hotgas flow path side 50 (radially inward side) of the shroud segment 30and by impingement cooling of the shroud segments 30 on the back side51.

Thermal barrier coatings (TBC) 31 are known in the art and may beapplied to the hot gas flow path side 50 of the shroud segments 30. Auseful TBC 31 may be Zircoat™ (by PRAXAIR) and may be applied by plasmaspray technique. The TBC 31 may provide thermal insulation against thehot gas flow and may reduce the cooling requirement of the shroudsegments 30. The turbine shroud assembly 25 may experience outwardradial movement due to thermal expansion. The forward end 72 of theshroud segment 30 and the aft end 73 of the shroud segment 30 mayexperience different rates of outward radial movement. The forward end72 may be the end of the shroud segment 30 that is closest to theforward hanger 32. The aft end 73 may be the end of the shroud segment30 that is closest to the aft hanger 33. The radial movement of theforward end 72 during engine operation may be a function of the thermalgrowth of the forward hanger 32. The radial movement of the aft end 73during engine operation may be a function of the thermal growth of theaft hanger 33. To compensate for the differential radial thermal growthbetween the forward hanger 32 and aft hanger 33, the thermal barriercoating 31 may be ground to a conical profile during manufacturing. Thethickness of the TBC 31 across the shroud segment 30 may vary from theforward end 72 to the aft end 73 such that the end of the shroud segment30 that experiences the greater outward radial movement may have thickerTBC 31. This may achieve an optimized cylindrical profile during engineoperation, which may maintain robust turbine tip clearance control.Because the present invention may be designed for applications that runat turbine inlet temperatures that can melt or fail even nickel basedsuperalloys, the shroud segments 30 may be sealed at their flatcircumferential surfaces 52 with ceramic spacer seals 38.

Ceramic spacer seals 38 may fit into the spacer channels 67, as shown inFIG. 5. One ceramic spacer seal 38 may fit simultaneously into two edgechannels 67 a and 67 b, one edge channel on each of two adjacentcircumferentially assembled shroud segments 30, as shown in FIG. 5. Oneceramic spacer seal 38 may be positioned in each spacer channel 67. Dueto manufacturing limitations and thermal growth considerations, theturbine shroud assembly 25 may comprise platform gaps 62 andfillet/chamfer gaps 63. The platform gap 62 may be the gap betweenadjacent shroud segments 30. The fillet/chamfer gap 63 may be the gapbetween a shroud segment 30 and a ceramic spacer seal 38. The ceramicspacer seals 38 may conform to the same circumferential cross section asthe shroud segments 30 but without the forward hooks 54 and aft hooks55, as shown in FIG. 3. The ceramic spacer seals 38 may comprise aceramic, such as monolithic silicon nitride ceramic. The ceramic spacerseals 38 may have a thickness between about 0.05 and about 0.4 inches.The thickness of the ceramic spacer seals 38 may be about equal to thedepth of the edge channels 67 a and 67 b.

The ceramic spacer seals 38 may be radially retained with forward spacerradial retainers 36 and aft spacer radial retainers 37, as shown in FIG.3. Any known retaining means may be useful for radially retaining theceramic spacer seals 38. The forward spacer radial retainers 36 and aftspacer radial retainers 37 may comprise known sheet metal retainingrings. During engine operation, the pressure loading of the turbineshroud assembly 25 may secure the ceramic spacer seals 38 to the shroudsegments 30, resulting in an exceptional high temperature seal thatwould not be possible utilizing the prior art. The shroud segments 30and ceramic spacer seals 38 may shelter the forward hanger rails 56 andaft hanger rail 58 from the hot flow path gases. The ceramic spacerseals 38 and the shroud segments 30 may be retained axially by theforward hanger 32 and the aft hanger 33.

The turbine shroud assembly 25 may comprise forward hanger 32 and afthanger 33. The forward hanger 32 and aft hanger 33 may comprise highstrength nickel base superalloys, such as Mar-M-247. For someapplications, the forward hanger 32 and aft hanger 33 may comprisemetal-alloys including nickel based alloys and cobalt based alloys.Useful metal alloys may include Single Crystal SC180 available fromHoneywell, Mar-M-247 Eqx material available from Martin Marietta, HA230available from Haynes International, and MA754 available from SpecialMetals. The forward hanger 32 and aft hanger 33 may be formed by knownmanufacturing methods. The forward hanger 32 and aft hanger 33 maygovern the overall thermal growth of the turbine shroud assembly 25. Thethermal expansion of the forward hanger 32 and aft hanger 33 may be afunction of hanger cooling and material selection. The forward hanger 32and aft hanger 33, which may be cooled independently, may allow foroptimal thermal growth and precise turbine tip clearance control. Forexample, the impingement cooling flow to the forward hanger 32 and afthanger 33 may be altered by altering the impingement cooling array 47,which in turn may alter the thermal expansion of the hangers and mayalter turbine tip clearance. The result may be improved turbine tipclearance and thus improved fuel consumption for the engine.

The turbine shroud assembly 25 may comprise an aft rope seal 35 betweenthe aft hanger rail 58 and the shroud segments 30. The aft rope seal 35may comprise a hybrid ceramic rope 61, as shown in FIG. 3. The hybridceramic rope 61 may comprise a rope core of braided ceramic fibersencapsulated in a metallic braided outer sheath. Useful hybrid ceramicropes 61 may include those similar to U.S. Pat. No. 5,301,595, which isincorporated herein by reference. The hybrid ceramic rope 61 may becompression tested to verify the load on the shroud segments 30 and maybe flow tested to quantify cooling flow as a function of compression.The diameter of a useful hybrid ceramic rope 61 may be between about0.04 and about 0.20 inches. The diameter of a preferred hybrid ceramicrope 61 may be between about 0.045 and about 0.080 inches. The diameterof a more preferred hybrid ceramic rope 61 may be between about 0.05 andabout 0.06 inches.

For proper positioning of the hybrid ceramic rope 61 and formation ofthe aft rope seal 35, the aft hanger rail 58 may be provided with anangled surface 58 a at the forward end of the radially inner surface, asshown in FIGS. 3 and 6. The angled surface 58 a may be provided by knownmanufacturing techniques. Force balance analysis may be utilized fordetermining a useful angle of the angled surface 58 a. The forcesaffecting the aft rope seal 35 may be at static equilibrium (forcebalance) during engine operation such that the hybrid ceramic rope 61may maintain contact with the shroud segment 30 and the angled surface58 a. A property of force is that the superposition of forces satisfiesthe laws of vector addition. When the forces acting on the hybridceramic rope 61 equal zero, the hybrid ceramic rope 61 will not be inmotion, thus not extruded. The sum of the forces affecting the aft ropeseal 35 may be zero. Factors affecting these forces may include thecompression of the hybrid ceramic rope 61, the pressure load from thecooling flow 60, the angle of the angled surface 58 a, and thefrictional coefficients of the materials. The magnitude of the forcesaffecting the aft rope seal 35 may vary by application. Friction testingsystems may be useful in determining the forces due to friction. Usefulfriction testing systems may include those described in U.S. Pat. No.5,377,525, which is incorporated herein by reference.

During manufacture of the turbine shroud assembly 25, an uncompressedform 64 of the hybrid ceramic rope 61 may be positionedcircumferentially about the shroud segments 30 and positioned radiallyinward of the aft hook 55 of the shroud segment 30. The aft hanger rail58 may be engaged axially to produce the compressed form 65, as seen inFIG. 8. During manufacture of the turbine shroud assembly 25, the hybridceramic rope 61 may be compressed between the aft hanger rail 58, theceramic spacer seals 38, and the shroud segments 30. In this compressedstate, the hybrid ceramic rope 61 may form an aft rope seal 35. As isknown in the art, a flow analysis test may be performed to determine thedesired compression. Useful flow analysis systems may include thosedescribed in U.S. Pat. No. 5,741,980, which is incorporated herein byreference.

The aft rope seal 35 may provide axial sealing and radial sealing. Thehybrid ceramic rope 61 may be compressed in such a way as tosimultaneously form a seal at a plurality of sealing areas 70, as shownin FIGS. 6 and 7. The sealing areas 70 may include a portion of the aftside of the ceramic spacer seal 38, a portion of the aft side of theshroud segment 30, the angled surface 58 a, the platform gap 62, and thefillet/chamfer gap 63. As can be seen, the hybrid ceramic rope 61 maycontact the aft face of the ceramic spacer seals 38 at a locationbetween adjacent shroud segments 30 to minimize radial leakage from theceramic spacer seals 38. The hybrid ceramic rope 61 may also seal thefillet/chamfer gap 63 created by the ceramic spacer seal 38 chamfer. Thereaction load for static equilibrium and final sealing on the angledsurface 58 a of the aft hanger rail 58 may occur to secure the aft ropeseal 35 and preclude extrusion of the hybrid ceramic rope 61 out of thegap between the angled surface 58 a and the shroud segments 30. The aftrope seal 35 may provide a robust flow circuit that minimizes hot gasingestion if a loss of back flow margin were to occur. For prior artseals, such as platform seals, when there is a loss of back flow margin,hot gas ingestion may occur unimpeded. The aft rope seal 35 of thepresent invention may be a physical impedance to hot gas ingestion.

The turbine shroud assembly 25 may comprise a forward rope seal 34between the forward hanger rails 56 and the shroud segments 30. Theforward rope seal 34 may comprise a hybrid ceramic rope 61. The forwardhanger rails 56 may be provided with o-ring grooves 56 a. An o-ringgroove 56 a may be positioned circumferentially on the radially innersurface of a forward hanger rail 56, as shown in FIG. 3. The o-ringgrooves 56 a may be formed by known methods and may be capable offorming a forward rope seal 34. The cross-sectional area of the o-ringgroove 56 a may be less than the cross-sectional area of the hybridceramic rope 61. The hybrid ceramic rope 61 may be positioned in theo-ring groove 56 a. The forward hanger rails 56 then may axially engagethe forward hook 54 of the shroud segment 30 to form the forward ropeseal 34.

During manufacturing of the turbine shroud assembly 25, the hybridceramic rope 61 may be positioned in the o-ring groove 56 a. The hybridceramic rope 61 may be compressed between the forward hanger rails 56and the shroud segments 30. In this compressed state, the hybrid ceramicrope 61 may produce a forward rope seal 34. The forward rope seal 34 maybe a typical “o-ring” type seal. The forward rope seal 34 may allowaxial movement between the shroud segments 30 and the forward hanger 32while maintaining a controlled amount of flow to purge the cavityforward of the forward hanger 32.

The turbine shroud assembly 25 may utilize the axial movement in theforward rope seal 34 to compensate for the differential axial thermalgrowth between the forward hanger 32 and aft hanger 33, while the aftrope seal 35 may maintain cooling back flow margin by controllingleakage across multiple sealing areas 70. Altering the compression ofthe hybrid ceramic rope 61, which forms the forward rope seal 34 and theaft rope seal 35, may alter the amount of cooling flow 60 exiting belowthe forward hanger 32 and aft hanger 33. A useful amount of cooling flow60 exiting below the forward hanger 32 and aft hanger 33 may vary byapplication. The rate of flow through the hybrid ceramic rope 61 may bea function of the compression of the hybrid ceramic rope 61. The hybridceramic rope 61 of the aft rope seal 35 may be more compressed than thehybrid ceramic rope 61 of the forward rope seal 34. This may allow agreater amount of cooling flow 60 to exit below the forward hanger 32than below the aft hanger 33.

Slots 77 may be positioned at the inner diameter of the forward hanger32 and extend radially outward, as seen in FIGS. 1 and 2. Slots 77 areknown in the art and may control thermal growth and stresses. Usefulslots 77 for reducing thermal expansion and stresses of the hanger mayinclude those described in U.S. Pat. No. 5,593,276, which isincorporated herein by reference. The useful dimensions and number ofslots 77 may be determined by known methods and may vary depending onapplication and hanger composition. For further control of thermalgrowth and for stress relief, stress relief openings 57 may bepositioned at the radially outer end of the slots 77 and extend axiallythrough the forward hanger 32, as shown in FIGS. 1–3. During engineoperation, cooling flow 60 may pass through the slots 77 and stressrelief openings 57 and may be channeled to help cool the forward hanger32 and provide purging of the hot flow path gas away from the turbineshroud assembly 25. Methods for producing the slots 77 and stress reliefopenings 57 are known in the art.

Useful methods for forming the slots 77 and stress relief openings 57may include electrical discharge machining (EDM). EDM applicationsrequire the use of a spark erosion machine. EDM applications are knownin the art and include drilling by spark erosion, which may be usefulfor forming the stress relief openings 57. Other known EDM applicationsinclude cutting by spark erosion using a flat electrode and cutting byspark erosion using a wire, both of which may be useful in forming theslots 77.

The turbine shroud assembly 25 may comprise a plenum assembly 40, asshown in FIG. 3. The plenum assembly 40 may be positioned between and incontact with the forward hanger 32 and aft hanger 33. The plenumassembly 40 may be retained between the forward hanger 32 and aft hanger33 by known retaining methods. In one embodiment, as shown in FIG. 3, abolting system 59 may retain the plenum assembly 40 between the forwardhanger 32 and the aft hanger 33. As shown in FIGS. 3 and 4, the plenumassembly 40 may comprise an axisymmetric plenum balloon 41, inletopenings 45, and flow metering openings 46. The axisymmetric plenumballoon 41 may have an impingement cooling array 47 there through, asshown in FIG. 4. Known manufacturing methods may be used to produce theplenum assembly 40. The plenum assembly 40 may comprise a sheet metalform and may be capable of receiving cooling flow 60.

During engine operation, a compressor (not shown) may be used to supplythe cooling flow 60. The cooling flow 60 may enter the plenum assembly40 at the inlet openings 45. The cooling flow 60 may pass through theinlet openings 45 and then enter the flow metering openings 46. Thecooling flow 60 may pass through the flow metering openings 46 and thenenter the axisymmetric plenum balloon 41. The-cooling flow 60 may thenexit the plenum assembly 40 through the impingement cooling array 47.

A vertical flange 39, shown in FIGS. 1–3, may be positioned radially outfrom and aft of the inlet openings 45 and may provide a robust method ofpressure recovery for axial flow combustor plenums. The cooling flow 60may be straightened as it flows radially inward through the inletopenings 45 and towards the flow metering openings 46. The inletopenings 45 may also serve to minimize the cooling flow circuitsensitivity to varying or uncertain orifice discharge coefficients thatare common with combustor plenums that suffer from significantcircumferential swirl. The straightening of the cooling flow 60 may bedue to the dimensions of the inlet openings 45. The length of the inletopenings 45 may be at least about two inlet hole diameters in length, asan example. The cross-sectional area of the inlet openings 45 may be atleast about 3 times the cross-sectional area of the flow meteringopenings 46. This may allow the flow metering openings 46 topredominantly control the overall flow to the axisymmetric plenumballoon 41. The plenum assembly may comprise at least one inlet opening45 and at least one flow metering opening 46. For every inlet opening 45there may be a corresponding flow metering opening 46 to provide auniform flow distribution to the impingement cooling array 47. In oneembodiment of the present invention the number of inlet openings 45 andcorresponding flow metering openings 46 may be about 36. The diameter ofthe inlet openings 45 may be between about 0.05 and about 0.20 inches.The diameter of the flow metering openings 46 may be between about 0.025and about 0.075 inches.

The axisymmetric plenum balloon 41 may comprise a sheet metal form andmay follow the contour of the back side 51 of the shroud segments 30.Impingement cooling effectiveness may be a function of impingementdistance. Impingement distance is the distance between the impingementopening and the surface to be cooled. The axisymmetric plenum balloon41, by following the contours of the shroud segments 30, may reduceimpingement distance and increase impingement cooling effectiveness.Known manufacturing methods may be used to form the axisymmetric plenumballoon 41. The impingement cooling array 47 may be machined into theaxisymmetric plenum balloon 41 and may comprise forward hanger coolingimpingement openings 42, aft hanger cooling impingement openings 43, andshroud cooling impingement openings 44.

During engine operation, a first portion 74 of cooling flow 60, as shownin FIG. 4, may pass through the forward hanger cooling impingementopenings 42 and may impinge the aft side of the forward hanger 32,thereby cooling the forward hanger 32. A second portion 75 of coolingflow 60 may pass through the aft hanger cooling impingement openings 43and may impinge the forward side of the aft hanger 33, thereby coolingthe aft hanger 33. A third portion 76 of cooling flow 60 may passthrough the shroud cooling impingement openings 44 and may impinge theback side 51 of the shroud segment 30, thereby cooling the shroudsegment 30 and the spacer seal 38. The heat transfer to the forwardhanger 32 and aft hanger 33, which govern the radial position of theshroud segments 30, which in turn govern the turbine tip clearance, maybe customized by altering the number or diameter of forward hangercooling impingement openings 42 and aft hanger cooling impingementopenings 43. This unique advantage over the prior art may help maintaintight turbine tip clearance. The impingement cooling array 47 may bepositioned and sized for optimal heat transfer with minimal cross flowdegradation. Any known method of heat transfer analysis may be useful.Factors affecting impingement cooling effectiveness may include the rateof flow of the cooling flow 60, cross flow degradation, impingementdistance, the diameters of the shroud cooling impingement openings 44,forward hanger cooling impingement openings 42, and aft hanger coolingimpingement openings 43. Cross flow degradation may be due to the spentair interfering with the intended impingement surfaces while the spentair exits the turbine shroud assembly 25.

The turbine shroud assembly 25 of the present invention may reduce theamount of cooling flow to the turbine shroud assembly 25 while enablingincreased turbine inlet temperatures above current production technologycapabilities.

As can be appreciated by those skilled in the art, the present inventionprovides improved turbine shroud assemblies and methods for theirproduction. A robust high temperature turbine shroud assembly isprovided that can operate in a higher temperature environment using lesscooling flow than the prior art. Also provided are turbine shroudassemblies having improved cooling efficiency. A turbine shroud assemblyhaving controlled cooling of the forward and aft hangers is alsoprovided. Further, a turbine shroud assembly capable of reducing turbinetip clearance is provided.

It should be understood, of course, that the foregoing relates topreferred embodiments of the invention and that modifications may bemade without departing from the spirit and scope of the invention as setforth in the following claims.

1. A turbine shroud assembly comprising: a plurality of shroud segmentsassembled circumferentially about a longitudinal engine centerline axis,at least one shroud segment having a forward hook and an aft hook; atleast one spacer channel positioned on the radially outward side of saidshroud segments, such that said at least one spacer channel is incontact with an interface of two said shroud segments; at least oneceramic spacer seal positioned in contact with said at least one spacerchannel, such that said at least one ceramic spacer seal is within saidat least one spacer channel; at least one forward hanger positionedradially outward from said shroud segments, said at least one forwardhanger having a forward hanger rail capable of engaging said forwardhook, said forward hanger rail having an o-ring groove positionedcircumferentially on a radially inward side; at least one aft hangerpositioned radially outward from said shroud segments, said at least oneaft hanger having an aft hanger rail capable of engaging said aft hook,said aft hanger rail having an angled surface positioned on the forwardedge of a radially inward side;and a plenum assembly positioned betweenand in contact with said at least one forward hanger and said at leastone aft hanger.
 2. The turbine shroud assembly of claim 1, furthercomprising a forward rope seal positioned between and in contact withsaid o-ring groove and said shroud segments.
 3. The turbine shroudassembly of claim 1, further comprising an aft rope seal positionedbetween and in contact with said angled surface and said shroudsegments.
 4. The turbine shroud assembly of claim 3, wherein said aftrope seal comprises a hybrid ceramic rope.
 5. The turbine shroudassembly of claim 1, wherein said shroud segments comprise a monolithicsilicon nitride ceramic.
 6. The turbine shroud assembly of claim 1,wherein said at least one forward hanger comprises a nickel based alloy.7. The turbine shroud assembly of claim 1, wherein said at least one afthanger comprises a nickel based alloy.
 8. The turbine shroud assembly ofclaim 1, wherein said at least one forward hanger has a plurality ofslots positioned at an inner diameter and extending radially outward. 9.The turbine shroud assembly of claim 1, wherein said at least oneceramic spacer seal comprises a monolithic silicon nitride ceramic. 10.The turbine shroud assembly of claim 1, further comprising a thermalbarrier coating (TBC) positioned on a radially inward side of saidshroud segments.
 11. A turbine shroud assembly comprising: a pluralityof ceramic shroud segments assembled circumferentially about alongitudinal engine centerline axis, each said ceramic shroud segmenthaving a forward hook and an aft hook;and a plurality of ceramic spacerseals positioned in contact with said ceramic shroud segments, such thateach one said ceramic spacer seal is in contact with the radiallyoutward side of two said ceramic shroud segments.
 12. The turbine shroudassembly of claim 11, further comprising a forward hanger radiallyoutward from and engaging said forward hook, and comprising an afthanger radially outward from and engaging said aft hook.
 13. The turbineshroud assembly of claim 12, further comprising a forward rope sealpositioned between said ceramic shroud segments and said forward hangersuch that said forward rope seal is radially inward from said forwardhanger.
 14. The turbine shroud assembly of claim 12, further comprisingan aft rope seal positioned between said ceramic shroud segments andsaid aft hanger such that said aft rope seal is radially inward fromsaid aft hanger.
 15. An apparatus for cooling turbine engine shroudsegments and hanger rails comprising: a plenum comprised of enclosingsurfaces of the segments and hanger rails; an axisymmetric plenumballoon positioned in the plenum and having an impingement cooling arraythere through; said impingement cooling array comprising holes with axesperpendicular to an outer surface of the axisymmetric plenum balloon; aplurality of flow metering openings in fluid communication with saidaxisymmetric plenum balloon; a plurality of inlet openings in flowcommunication with said flow metering openings; and wherein an outersurface of the axisymmetric plenum balloon follows a contour of aninternal surface of the plenum so that each of the holes of theimpingement array are oriented to provide flow in a direction that isperpendicular to a corresponding underlying portion of the internalsurface of the plenum.
 16. The apparatus of claim 15, wherein saidimpingement cooling array comprises a plurality of forward hangercooling impingement openings, a plurality of aft hanger coolingimpingement openings, and a plurality of shroud cooling impingementopenings.
 17. The apparatus of claim 15, wherein a cross-sectional areaof said inlet openings is at least about three times a cross-sectionalarea of said flow metering openings and wherein the inlet openings areaxially aligned with the flow metering openings.
 18. The apparatus ofclaim 17, further comprising a vertical flange positioned radiallyoutward from and aft of said inlet openings.
 19. The apparatus of claim15, wherein said axisymmetric plenum balloon comprises a sheet metalform.
 20. A rope seal apparatus for use between a turbine shroud andhangers comprising: first and a second compressed hybrid ceramic ropespositioned between and in contact with said turbine shroud and saidhangers, such that said hangers are radially outward from saidcompressed hybrid ceramic ropes; the first hybrid ceramic rope beingadjacent the forward hanger; the second hybrid ceramic rope beingadjacent the aft hanger; the second hybrid ceramic rope being compressedto a greater degree than the first hybrid ceramic rope.
 21. The ropeseal apparatus of claim 20, wherein said turbine hanger comprises aforward hanger having a circumferential o-ring groove positioned on aradially inward side, and said compressed hybrid ceramic rope is incontact with said circumferential o-ring groove.
 22. The rope sealapparatus of claim 20, wherein said turbine hanger comprises an afthanger having a circumferential angled surface positioned on a forwardedge of the radially inward side, and said compressed hybrid ceramicrope is in contact with said angled surface.
 23. The rope seal apparatusof claim 20, wherein said turbine shroud comprises a plurality ofceramic shroud segments.
 24. A method of shielding a turbine engine froma hot gas flow path there through comprising the steps of: providing aplurality of ceramic shroud segments assembled circumferentially about alongitudinal engine centerline axis through said hot gas flow path, aforward hanger radially outward from and forward of said ceramic shroudsegments, an aft hanger radially outward from and aft of said ceramicshroud segments, and a plenum assembly between and in contact with saidforward hanger and said aft hanger;and supplying a cooling flow to saidplenum assembly such that said cooling flow impinges said ceramic shroudsegments, said forward hanger, and said aft hanger.
 25. The method ofclaim 24, further comprising a step of positioning a hybrid rope sealbetween said forward hanger and said ceramic shroud segments.
 26. Themethod of claim 24, further comprising a step of positioning a hybridrope seal between said aft hanger and said ceramic shroud segments. 27.The method of claim 24, further comprising a step of positioning aceramic spacer seal in contact with a radially outward side of eachinterface of two ceramic shroud segments.
 28. The method of claim 27,wherein said ceramic spacer seals comprise a monolithic silicon nitrideceramic.
 29. The method of claim 24, wherein said plenum assemblycomprises an axisymmetric plenum balloon having an impingement coolingarray there through, a plurality of flow metering openings in fluidcommunication with said axisymmetric plenum balloon, and a plurality ofinlet openings in flow communication with said flow metering openings.